1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to turbine rotor blade with serpentine flow cooling.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
A gas turbine engine, such as a large frame heavy duty industrial gas turbine (IGT) engine, includes a turbine with one or more rows of stator vanes and rotor blades that react with a hot gas stream from a combustor to produce mechanical work. The stator vanes guide the hot gas stream into the adjacent and downstream row of rotor blades. The first stage vanes and blades are exposed to the highest gas stream temperatures and therefore require the most amount of cooling.
The efficiency of the engine can be increased by using a higher turbine inlet temperature. However, increasing the temperature requires better cooling of the airfoils or improved materials that can withstand these higher temperatures. Turbine airfoils (vanes and blades) are cooled using a combination of convection and impingement cooling within the airfoils and film cooling on the external airfoil surfaces.
FIG. 1 shows a blade external heat transfer coefficient (HTC) profile for a first stage turbine rotor blade in an industrial gas turbine engine. As seen in FIG. 1, the airfoil leading edge and trailing edge as well as a forward region of the suction side surface experiences high hot gas heat transfer coefficient while the mid-chord section of the airfoil is at a lower hot gas HTC. Thus, the hottest parts of the first stage blade are on the leading and trailing edges and on the suction side wall just downstream from the leading edge region.
FIG. 2 shows a prior art turbine rotor blade with a 5-pass serpentine flow aft flowing cooling circuit, FIG. 3 shows a cross section view along a radial line of the FIG. 2 blade cooling circuit and FIG. 4 shows a flow diagram for the blade cooling circuit for the FIG. 2 blade. The first leg of the 5-pass serpentine circuit is located at the leading edge to provide cooling for this section. The last and fifth leg is located along the trailing edge region and is connected to a row of trailing edge exit slots that provide cooling for the trailing edge region of the blade. No film cooling holes are sued in the FIG. 2 blade and therefore all of the cooling air that flows into the first leg eventually flows into the last leg to be discharged out through the exit slots.
One major problem with the FIG. 1 design is that the fresh cooling air passing through the first leg is heated and then passed through the next three legs in the airfoil mid-chord region before passing along the last leg in the trailing edge region. Thus, the cooling air to be used in the trailing edge region is heated up more than necessary and the airfoil mid-chord region is over-cooled because the cooling air from the first leg passes through the mid-chord region before passing through the trailing edge region. The over-heated cooling air used for the T/E region will induce hot spots in the T/E metal temperature which will cause erosion damage and thus a shortened blade life, especially for an engine like an IGT engine that requires continuous operating periods of over 40,000 hours before shutdown.
In order to over-come some of the over-heating of the T/E region and over-cooling of the airfoil mid-chord region described in the FIG. 1 blade, a redistribution of cooling air within the 5-pass serpentine flow cooling circuit is required. U.S. Pat. No. 6,139,269 issued to Liang on Oct. 31, 2000 and entitled TURBINE BLADE WITH MULTI-PASS COOLING AND COOLING AIR ADDITION discloses a blade with a 5-pass forward flowing serpentine cooling circuit with cooling air addition in turns between the second and third legs and between the fourth and fifth legs to resupply the serpentine circuit with cooler fresh cooling air in the airfoil mid-chord region.
U.S. Pat. No. 6,340,047 issued to Frey on Jan. 22, 2002 and entitled CORE TIED CAST AIRFOIL discloses a blade with a 5-pass aft flowing serpentine flow cooling circuit in which fresh cooling air from the root is injected into the turns between the second and third legs and between the fourth and fifth legs through ball braze holes. U.S. Pat. No. 6,966,756 issued to McGrath et al. on Nov. 22, 2005 and entitled TURBINE BUCKET COOLING PASSAGES AND INTERNAL CORE FOR PRODUCING THE PASSAGES and U.S. Pat. No. 7,674,093 issued to Lee et al on Mar. 9, 2010 and entitled CLUSTER BRIDGED CASTING CORE discloses similar fresh cooling air resupply passages for a serpentine flow cooling circuit within a blade that use ball braze holes to close out the ceramic core support holes.